Technical Field of Invention
The invention disclosed and taught herein relates generally to a method and system of automatically identifying and characterizing composite laminate structures, or laminates, using ultrasonic non-destructive testing (NDT) techniques. More specifically, the invention disclosed herein relates to a method and system of automatically detecting each layer, or ply, of material in a laminate and determining the bulk properties of the laminate based on the properties of the constituent plies in order to generate a failure envelope for the laminate. The invention disclosed and taught herein also relates to a method and system of simulating an ultrasonic scan of the individual plies of a laminate.
Description of Related Art
Composite laminates, or laminates, are typically composed of individual layers of materials that have directionally dependent material properties. Each layer is commonly known as a lamina or ply, and the plies are combined in layers to create a bulk structure that forms the laminate. Knowledge of the individual lamina configuration is important because of the significant effect each lamina has on the final properties of the laminate. For example, in unidirectional fiber orientation plies, the ply is considerably stronger in the fiber orientation direction than in any other direction. The choice of orientation, thickness, stacking sequence, or other property of a lamina within the final composite, will drastically alter the final processed material properties of the laminate.
Composite laminates are used extensively in a variety of structural applications and in numerous industries. For example, carbon fiber reinforced polymers/plastics (CFRP) are a commonly-used type of composite laminates in the aerospace, automotive and other industries. Although CFRPs can be relatively expensive, their superior strength-to-weight ratios make them more desirable than other types of materials. This is reflected by the widespread and steadily increasing use of CFRP components in fixed and rotary wing airframes, for example. Alternatively, a large industry exists that implements alternative reinforcements sacrificing the ultimate tensile strength for other design parameters such as cost, processing ease, etc. Commonly employed fibers may include, but are not limited to, fiberglass, Kevlar, aramid, and other synthetic fibers, as well as a wide variety of natural fibers used as fillers.
Manufacturing carbon fibers usually involves a process where a single continuous carbon fiber filament is constructed with a diameter of roughly 0.005 mm to 0.010 mm. For the type of high quality products used in aerospace applications, a typical fiber diameter will be on the order of 0.005 mm. These filaments are 93% to 95% carbon and have a linear mass of roughly 6.6 grams per meter (g/m). Individual filaments are then wound into a “tow” (i.e., thread or ribbon) that may then be used for various applications. Typical tows have between 3,000 and 12,000 filaments depending on the product application. A 3,000-filament tow has a linear mass of about 0.2 g/m and can be between 0.375 mm and 1.5 mm wide and between 0.2 mm and 0.05 mm thick. By comparison, the diameter of an average piece of thread is approximately 0.375 mm for a 3,000-filament thread.
The tows may be woven into a pattern and then impregnated with resin to form an individual lamina that may then be stacked on other laminas to create a composite laminate layup. The main geometries for an individual lamina are: Percent Warp=percent of orthogonal fibers by weight (where 0% means fibers are unidirectional, and 50% means fibers are woven); Areal Density=g/m2 of fiber in a given lamina; Thread Count=number of individual fiber threads in an individual tow; Tow Width=width in mm of an individual tow; Layer Thickness=thickness of an individual lamina in mm. In contrast, the main geometries for a completed composite layup are: number of laminas, fiber orientation of individual lamina, the lamina type (i.e., woven versus non-woven, woven type, material makeup), and layup method of individual layers.
The need to repair and modify laminated composites has stretched the capability of existing non-destructive inspection (NDI) techniques. Specifically, to properly modify or repair laminated composites, sufficient fidelity of the underlying microstructure of the composite plies is required to understand the baseline (i.e., unmodified, unrepaired) structure, identify and quantify any as-installed modifications, and analyze the as-installed components for FAA, automotive, and other industry certifications.
In addition, the manufacturing process for composite laminates, as for other materials, inherently includes some variability that can affect the performance of the final part. As such, it is desirable to account for these manufacturing uncertainties and tolerances when quantifying the expected structural response of composite laminates. It is also desirable to quantify the impact of manufacturing defects and varying material properties on a composite laminate's performance. Conversely, if the configuration of a composite laminate can be determined within some degree of confidence, it would be desirable to quantify the expected structural response and life expectancy of that laminate.
Having the ability to detect manufacturing, installation, or usage effects on a composite laminate, with resolutions on the order of individual lamina dimensions and as a function of affected lamina layer, may also minimize modification and maintenance design conservatism or safety margins, leading to reduced manufacturing, installation, and test costs.
The ability to quantify a composite laminate's expected structural response would be particularly useful where the manufacture's data and information about the composite laminate are limited or perhaps unavailable. The problem is compounded by the need in many instances to ascertain the composite laminate's as-fabricated structural characteristics, including stiffness, failure envelope, and the like, without performing destructive testing.
Accordingly, what is needed is a system and method for identifying and characterizing a composite laminate's internal structure using ultrasonic NDT techniques. More particularly, what is needed is a system and method for quantifying the composite laminate's expected structural response based on the assessed properties of the individual laminas.